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Artykuły w czasopismach na temat „Thrust specific impulse”

Autor: Grafiati

Data publikacji: 2 czerwca 2025

Data aktualizacji: 31 lipca 2025

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1

Navalino, RDA, NRS Muda, MAE Hafizah, and Y. Ruyat. "Analysis of carbon nano particle variant as the propellant fuel to increase specific impulses of rockets." F1000Research 12 (May 29, 2024): 1414. http://dx.doi.org/10.12688/f1000research.138276.3.

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Background This study compares propellant fuels’ specific thrust and impulse parameters using nanocarbon variant fuels from coconut shells and coal. Specific impulses and impulses are essential parameters that determine rocket performance. The specific thrust and impulses are influenced by fuel type, material composition, heat flow, and burning time parameters. The characteristics of nanocarbon as a fuel are proven to have high-value features and a long combustion time. This parameter is essential to add value to specific thrusts and impulses. Methods The method to prove the quality of solid f

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2

Navalino, RDA, NRS Muda, MAE Hafizah, and Y. Ruyat. "Analysis of carbon nano particle variant as the propellant fuel to increase specific impulses of rockets." F1000Research 12 (October 26, 2023): 1414. http://dx.doi.org/10.12688/f1000research.138276.1.

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Background: This study compares propellant fuels' specific thrust and impulse parameters using nanocarbon variant fuels from coconut shells and coal. Specific impulses and impulses are essential parameters that determine rocket performance. The specific thrust and impulses are influenced by fuel type, material composition, heat flow, and burning time parameters. The characteristics of nanocarbon as a fuel are proven to have high-value features and a long combustion time. This parameter is essential to add value to specific thrusts and impulses. Methods: The method to prove the quality of solid

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3

Navalino, RDA, NRS Muda, MAE Hafizah, and Y. Ruyat. "Analysis of carbon nano particle variant as the propellant fuel to increase specific impulses of rockets." F1000Research 12 (December 11, 2023): 1414. http://dx.doi.org/10.12688/f1000research.138276.2.

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Background: This study compares propellant fuels’ specific thrust and impulse parameters using nanocarbon variant fuels from coconut shells and coal. Specific impulses and impulses are essential parameters that determine rocket performance. The specific thrust and impulses are influenced by fuel type, material composition, heat flow, and burning time parameters. The characteristics of nanocarbon as a fuel are proven to have high-value features and a long combustion time. This parameter is essential to add value to specific thrusts and impulses. Methods: The method to prove the quality of solid

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4

K., F. Oyedeko*1 &. A. Onyieagho2. "EFFECT OF PROPELLANT FORMULATION ON PROPELLANT PROPERTIES." INTERNATIONAL JOURNAL OF ENGINEERING SCIENCES & RESEARCH TECHNOLOGY 7, no. 8 (2018): 305–13. https://doi.org/10.5281/zenodo.1345623.

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This paper present the effect of propellant formulations on propellant properties is very crucial to the quality of propellant used in rocket motor. As poor propellant formulation usually leads to unstable combustion and this can lead to a total destruction of the entire structure of the solid rocket motor. For this research work, the centre compositions selected were based on calculation for percentage mass composition from the stoichiometric equation for the combustion process. Hollow bates grains used for the design consist of a mass of propellant of 6 kg, core diameter of 0.050m, throat di

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5

Berliand, M., and R. Ishchenko. "CALCULATION OF THE EXPANSION PROCESS OF THE FLOW OF WORKING BODY IN THE NOZZLE OF A LIQUID-PHASE NUCLEAR ROCKET ENGINE." Slovak international scientific journal, no. 95 (May 15, 2025): 86–92. https://doi.org/10.5281/zenodo.15427667.

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In this work, flow parameters of the working body (hydrogen) were calculated at the inlet, throat, and exit sections of the liquid-phase nuclear rocket engine nozzle. A method was proposed for determining the specific impulse of thrust while accounting for losses due to scattering, friction, and variations in the specific enthalpy of the working body caused by non-adiabatic effects. The calculated specific impulse exceeded 9500 m/s, significantly surpassing the performance of chemical rocket engines. The evaluation of the thermogasdynamic parameters of the working body expansion confirmed the

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6

Cook, Ronald, James A. Nabity, and John W. Daily. "Characterizing Propellants for Variable-Thrust/Specific Impulse Colloid Thrusters." Journal of Propulsion and Power 33, no. 6 (2017): 1325–31. http://dx.doi.org/10.2514/1.b36495.

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Kammash, Terry, Myoung-Jae Lee, and David I. Poston. "High-Thrust-High-Specific Impulse Gasdynamic Fusion Propulsion System." Journal of Propulsion and Power 13, no. 3 (1997): 421–27. http://dx.doi.org/10.2514/2.5180.

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8

Taheri, Ehsan, and John L. Junkins. "How Many Impulses Redux." Journal of the Astronautical Sciences 67, no. 2 (2019): 257–334. http://dx.doi.org/10.1007/s40295-019-00203-1.

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AbstractA central problem in orbit transfer optimization is to determine the number, time, direction, and magnitude of velocity impulses that minimize the total impulse. This problem was posed in 1967 by T. N. Edelbaum, and while notable advances have been made, a rigorous means to answer Edelbaum’s question for multiple-revolution maneuvers has remained elusive for over five decades. We revisit Edelbaum’s question by taking a bottom-up approach to generate a minimum-fuel switching surface. Sweeping through time profiles of the minimum-fuel switching function for increasing admissible thrust m

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9

Cojocea, Andrei Vlad, Ionuț Porumbel, Mihnea Gall, and Tudor Cuciuc. "Experimental Thrust and Specific Impulse Analysis of Pulsed Detonation Combustor." Applied Sciences 14, no. 14 (2024): 5999. http://dx.doi.org/10.3390/app14145999.

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Detonation combustion represents a significant advancement in efficiency over traditional deflagration methods. This paper presents a Pulsed Detonation Combustor (PDC) model that is designed with an aerodynamic mixing chamber featuring Hartmann–Sprenger resonators and crossflow injection. This design enhances operational cycle frequency and enables sustained detonation over short distances (below 200 mm). The PDC’s performance was evaluated through a comprehensive full-factorial experimental campaign, incorporating four factors with four discrete levels each. Testing was conducted using both h

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10

DHAMA, Sanjeev Kumar, T. K. JINDAL, and S. K. MANGAL. "Influence of Nozzle Type, Divergence Angle and Area Ratio on Impulse of Pulse Detonation Engine." INCAS BULLETIN 12, no. 2 (2020): 35–45. http://dx.doi.org/10.13111/2066-8201.2020.12.2.4.

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The influence of nozzle geometry on the impulse produced by the single cycle Pulse detonation engine (PDE) was experimentally investigated. For each experiment the nozzles were attached at the end of the engine. The impulse produced by the pulse detonation engine was calculated from the measured thrust. The thrust measurement was done by sliding the engine on the central bar of the thrust stand. The main structure of the basic PDE has a detonation tube with one terminal closed, a Schelkin spiral used as deflagration to detonation device, and a thrust stand to support the structure. Stoichiomet

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11

Lin, Zhen, and Yong Li. "Performance Analysis of Bipropellant Propulsion System on Special Working Conditions." Applied Mechanics and Materials 390 (August 2013): 296–300. http://dx.doi.org/10.4028/www.scientific.net/amm.390.296.

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At the end of a test satellite, oxidant of the propulsion system was exhausted ahead of fuel. Long-term monopropellant working condition without combustion, cool gas flow working condition and interim working condition were experienced, which trouble the de-orbit control of the satellite. In order to evaluate the performance of the propulsion system on the special working conditions during the satellite de-orbit, theoretic analysis of the specific thrusters is introduced. And effective calculation method is built. Based on the method and obtained remote data, the thrust and specific impulse of

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12

Jobanpreet Singh. "Performance and Combustion Analysis of Solid Rocket Propellant Using Aluminum Powder, Ammonium Perchlorate, and HTPB." International Journal of Advanced Research and Interdisciplinary Scientific Endeavours 2, no. 3 (2025): 519–28. https://doi.org/10.61359/11.2206-2513.

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In this paper, the design and performance evaluation of a small-scale solid propellant rocket were discussed to be made up of a mixture of powdered aluminum (Al), ammonium perchlorate (NH₄ClO₄), and hydroxyl-terminated polybutadiene (HTPB). A 92.5 cm with a diameter of 5 cm and total mass of 0.854 kg, well-designed rocket using Creo software with optimal structural and aerodynamic performance was considered. Parametric parameters include a center of gravity located at 69.4 cm with a center of pressure of 61.8 cm, therefore developing excellent stability in the case of flight. The propellant fo

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13

Cheng, Dah Yu. "Deflagration Plasma Thruster, a High Thrust, High Specific Impulse Device for Nuclear Space Propulsion." Fusion Technology 20, no. 4P2 (1991): 730–34. http://dx.doi.org/10.13182/fst91-a11946928.

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14

Zhong, Yubin, Fabrizio Ponti, Francesco Barato, et al. "Design and Analysis of a Micro–Electro-Mechanical System Thruster for Small Satellites and Low-Thrust Propulsion." Aerospace 12, no. 3 (2025): 172. https://doi.org/10.3390/aerospace12030172.

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As a cost-effective and versatile solution, small satellites are increasingly being considered for space exploration. However, one of the major challenges in deploying small satellites for high total impulse missions, particularly deep space exploration, lies in the propulsion system. These missions face strict constraints in terms of volume, mass, and power budgets. This paper proposes a potential solution to this issue through the design of a bipropellant MEMS thruster. Simulation results indicate that this type of thruster offers superior performance compared to the monopropellant propulsio

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15

Palla, Daniele, and Gabriele Cristoforetti. "Laser–Accelerated Plasma–Propulsion System." Applied Sciences 11, no. 21 (2021): 10154. http://dx.doi.org/10.3390/app112110154.

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In this paper, the laser-accelerated plasma–propulsion system (LAPPS) for a spacecraft is revisited. Starting from the general properties of relativistic propellants, the relations between specific impulse, engine thrust and rocket dynamics have been obtained. The specific impulse is defined in terms of the relativistic velocity of the propellant using the Walter’s parameterization, which is a suitable and general formalism for closed–cycle engines. Finally, the laser-driven acceleration of light ions via Target Normal Sheath Acceleration (TNSA) is discussed as a thruster. We find that LAPPS i

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16

K., F. Oyedeko*1 &. A. Onyieagho2. "OPTIMISATION OF THE FORMULATION OF A DOUBLE – BASED SOLID PROPELLANT." INTERNATIONAL JOURNAL OF ENGINEERING SCIENCES & RESEARCH TECHNOLOGY 7, no. 8 (2018): 296–304. https://doi.org/10.5281/zenodo.1345619.

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The study investigated the optimization of the formulation of a double – based solid propellant using Sorbitol as fuel and Potassium nitrate as oxidizer. This research provides an insight into a way of preventing the negative effect of poor propellant formulation on the predetermined rocket mission. The design is based on the fact that specific impulse, temperature, density and thrust of the product are functionally related to specific propellant formulation and these are fitted to multiple regression equations describing responses to optimal formulation using response surface methodolog

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17

Pakhomov, Andrew V., M. Shane Thompson, Wesley Swift, and Don A. Gregory. "Ablative Laser Propulsion: Specific Impulse and Thrust Derived from Force Measurements." AIAA Journal 40, no. 11 (2002): 2305–11. http://dx.doi.org/10.2514/2.1567.

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18

Ketsdever, Andrew D., Brian C. D'Souza, and Riki H. Lee. "Thrust Stand Micromass Balance for the Direct Measurement of Specific Impulse." Journal of Propulsion and Power 24, no. 6 (2008): 1376–81. http://dx.doi.org/10.2514/1.36921.

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19

Pakhomov, A. V., M. S. Thompson, W. Swift, and D. A. Gregory. "Ablative laser propulsion - Specific impulse and thrust derived from force measurements." AIAA Journal 40 (January 2002): 2305–11. http://dx.doi.org/10.2514/3.15323.

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20

Zheng, Jiaxuan, Beinan Jia, and Yongjun Jian. "Steric Effects on Space Electroosmotic Thrusters in Soft Nanochannels." Mathematics 9, no. 16 (2021): 1916. http://dx.doi.org/10.3390/math9161916.

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The influence of steric effects on the performances of space electroosmotic thrusters (EOTs) was numerically delineated in soft nanochannels for which its walls are covered with polyelectrolyte materials. The size effect of the ionic species, namely the steric effect, is neglected in many previous research studies, but it has vital influences on electrostatic potential and electroosmotic velocity, which is further introduced into the present study in order to understand and improve the exploration of nano electroosmotic thrusters with soft channels. The thruster’s thrust, specific impulse, tot

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21

Sliusariev, V. V., and I. V. Bilotserkovsky. "ANALYSIS OF THE CORRECTNESS APPROACHES TO DETERMINING THE SPECIFIC IMPULSE OF THRUST GENERATED BY THE WORKING FLUID OF A GAS GENERATOR CYCLE ROCKET ENGINE TURBINE." System design and analysis of aerospace technique characteristics 34, no. 1 (2024): 93–106. http://dx.doi.org/10.15421/472408.

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For liquid rocket engines (LREs) utilizing the gas generator cycle, the resulting engine thrust is composed of two components: the thrust of the engine chamber and the thrust of the turbine exhaust nozzle. Over the years of LRE utilization, methodologies for calculating ideal thermo gas dynamic parameters of the chamber have been developed, large volumes of statistical data regarding specific impulse losses in engine chambers have been obtained, and approaches for predicting specific impulse and, consequently, thrust of the main chambers have been refined. A much more complex situation arises

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22

Underwood, Thomas C., William M. Riedel, and Mark A. Cappelli. "Dual mode operation of a hydromagnetic plasma thruster to achieve tunable thrust and specific impulse." Journal of Applied Physics 130, no. 13 (2021): 133301. http://dx.doi.org/10.1063/5.0051467.

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23

Kirshina, A. A., A. A. Levikhin, and A. Yu Kirshin. "Comparative results of computational and theoretical study of the annular nozzle with a flat central body." VESTNIK of Samara University. Aerospace and Mechanical Engineering 23, no. 2 (2024): 28–39. http://dx.doi.org/10.18287/2541-7533-2024-23-2-28-39.

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One of the ways to improve specific characteristics of the power plant of a launch vehicle for ladeploying payload to the near-Earth space is to provide the possibility of operation of a fixed nozzle in the design mode over the whole active leg of the flight trajectory. The nozzle should be compact, lightweight, well-cooled. For detailed testing of the possibility of introducing a nozzle into the rocket engine chamber it is necessary to be able to quickly assess the true value of the thrust and the specific impulse the chamber with such a nozzle can achieve. This article presents the results o

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24

Olena, Kositsyna, Varlan Kostiantyn, Dron Mykola, and Kulyk Oleksii. "Determining energetic characteristics and selecting environmentally friendly components for solid rocket propellants at the early stages of design." Eastern-European Journal of Enterprise Technologies 6, no. 6 (114) (2021): 6–14. https://doi.org/10.15587/1729-4061.2021.247233.

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This paper has investigated the possibility to theoretically calculate a value of the specific impulse for highly energetic compositions using only two parameters – the heat of the reaction and the number of moles of gaseous decomposition reaction products. Specific impulse is one of the most important energetic characteristics of rocket propellant. It demonstrates the level of achieving the value of engine thrust and propellant utilization efficiency. Determining the specific impulse experimentally is a complex task that requires meeting special conditions. For the stage of synthesis of

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25

Zolotko, O. E., O. V. Zolotko, O. V. Sosnovska, O. S. Aksyonov, and I. S. Savchenko. "Tte stage deorbiting with a deceleration pulse detonation engine." Kosmìčna nauka ì tehnologìâ 27, no. 4 (2021): 32–41. http://dx.doi.org/10.15407/knit2021.04.032.

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The article discusses the issues related to reducing the amount of space debris from rocket stages. The main ways to remove the separable part of a rocket from a space orbit are: the usе of a deceleration detonation propulsion system; gasification of fuel residues and the use of a gas-reactive deceleration pulse system; continuation of the work of the main propulsion system after the separation of stages; the use of a harpoon to capture the rocket stage and the use of sail for its further braking; the use of anti-missile or combat lasers to destroy a stage on the orbit followed by the stage fr

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26

Zheng, Jiaxuan, Siyi An, and Yongjun Jian. "Steric Effects on Electroosmotic Nano-Thrusters under High Zeta Potentials." Mathematics 9, no. 24 (2021): 3222. http://dx.doi.org/10.3390/math9243222.

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Here, space electroosmotic thrusters in a rigid nanochannel with high wall zeta potentials are investigated numerically, for the first time, considering the effect of finite size of the ionic species. The effect, which is called a steric effect, is often neglected in research about micro/nano thrusters. However, it has vital influences on the electric potential and flow velocity in electric double layers, so that the thruster performances generated by the fluid motion are further affected. These performances, including thrust, specific impulse, thruster efficiency, and the thrust-to-power rati

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27

Senent, Juan, Cesar Ocampo, and Antonio Capella. "Low-Thrust Variable-Specific-Impulse Transfers and Guidance to Unstable Periodic Orbits." Journal of Guidance, Control, and Dynamics 28, no. 2 (2005): 280–90. http://dx.doi.org/10.2514/1.6398.

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Mendoza-Anchondo, Reyna Judith, Cornelio Alvarez-Herrera, and José Guadalupe Murillo-Ramírez. "Visualization and Parameters Determination of Supersonic Flows in Convergent-Divergent Micro-Nozzles Using Schlieren Z-Type Technique and Fluid Mechanics." Fluids 10, no. 2 (2025): 40. https://doi.org/10.3390/fluids10020040.

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Small-scale and supersonic convergent-divergent type micro-nozzles with characteristic sizes of around a few centimeters and exit and throat radii of tenths of millimeters were the subjects of this study. Using the schlieren Z-type optical technique, the supersonic airflows established at the exit of seven nozzles were visualized. The dependence of the shock cell characteristics on the nozzle pressure ratio (NPR), defined as the ratio of stagnation pressure to atmospheric pressure, was analyzed. The dependence of the nozzle thrust and the specific impulse on the NPR ratio and the mass flow rat

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29

Tahsini, Amir Mahdi. "Numerical Prediction of the Magneto Plasma Dynamic Thrusters’ Performance." Applied Mechanics and Materials 598 (July 2014): 239–43. http://dx.doi.org/10.4028/www.scientific.net/amm.598.239.

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The performance of the magnetoplasmadynamic thruster is predicted using numerical simulation. The thruster mode is self-induced magnetic field with cylindrical electrodes. The dependence of the thrust level, specific impulse, and the mass flow rate in different total electric currents is investigated. The AUSM+ scheme is utilized to develop a numerical procedure and the accurate method is used to simulate the propellant injection rate. Besides the performance curves prediction, the results show the importance of the effect of inlet modeling on the thruster’s actual specific impulse.

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Radhakrishnan, Kanmaniraja, Dong Hwi Ha, and Hyoung Jin Lee. "Effect of Multicoaxial Injectors on Nitrogen Film Cooling in a GCH4/GO2 Thrust Chamber for Small-Scale Methane Rocket Engines: A CFD Study." Aerospace 11, no. 9 (2024): 744. http://dx.doi.org/10.3390/aerospace11090744.

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Improper film cooling design and positioning of an injector in the face plate cause thermal damage to the thrust chamber wall and lead to rocket engine failures. An experimental combustor with five shear coaxial injectors was damaged owing to inadequate film cooling distribution on the thrust chamber wall. The present study aimed to simulate the experimental test case and investigate the causes of the thermal damage. In the simulation, gaseous methane and oxygen were injected at the inner and outer inlets of the shear coaxial injectors and nitrogen, used as the coolant, was injected near the u

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31

Takao, Yoshinori, Takeshi Takahashi, Koji Eriguchi, and Kouichi Ono. "Microplasma thruster for ultra-small satellites: Plasma chemical and aerodynamical aspects." Pure and Applied Chemistry 80, no. 9 (2008): 2013–23. http://dx.doi.org/10.1351/pac200880092013.

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A microplasma thruster has been developed of electrothermal type using azimuthally symmetric microwave-excited microplasmas. The microplasma source was ~2 mm in diameter and ~10 mm long, being operated at around atmospheric pressures; the micronozzle was a converging-diverging type, having a throat ~0.2 mm in diameter and ~1 mm long. Numerical and experimental results with Ar as a working gas demonstrated that this miniature electrothermal thruster gives a thrust of >1 mN, a specific impulse of ~100 s, and a thrust efficiency of ~10 % at a microwave power of <10 W, making it applicable t

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32

Jia, FeiDa, Dong Qiao, HongWei Han, and XiangYu Li. "Efficient optimization method for variable-specific-impulse low-thrust trajectories with shutdown constraint." Science China Technological Sciences 65, no. 3 (2022): 581–94. http://dx.doi.org/10.1007/s11431-021-1949-0.

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Хмелевской, И. А., та Д. А. Томилин. "Исследование режимов горения разряда "спица" и "колокол" на холловском двигателе мощностью 1.5 kW". Письма в журнал технической физики 46, № 10 (2020): 31. http://dx.doi.org/10.21883/pjtf.2020.10.49429.18240.

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Two stable operating modes, which were called “jet” and “bell” mode, were investigated for Hall thruster with power 1.5 kW. Performances (thrust, thrust specific impulse) in two modes were investigated. The study were conducted for discharge voltage 300-800 V and gas flow rate from 1.5 to 3.0 mg/s. This result is compared with the previous study, which carried out on geometrically similar but bigger Hall thruster model.

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Wei, Shih-Sin, Jui-Cheng Hsu, Hsi-Yu Tso, and Jong-Shinn Wu. "Investigation of Performance Stability of a Nytrox Hybrid Rocket Propulsion System." Aerospace 12, no. 5 (2025): 372. https://doi.org/10.3390/aerospace12050372.

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Nitrous oxide is a highly suitable oxidizer for hybrid rockets due to its self-pressurizing properties, moderate cost, and high accessibility. However, its vapor pressure and density are highly dependent on ambient temperature, requiring careful consideration of temperature variations in real applications. To mitigate this issue, an oxidizer called Nytrox was produced by adding a small fraction of oxygen to bulk nitrous oxide. This modification enables the hybrid rocket propulsion system to maintain a nearly constant average thrust and total impulse across a wide range of ambient temperatures.

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Kositsyna, Olena, Kostiantyn Varlan, Mykola Dron, and Oleksii Kulyk. "Determining energetic characteristics and selecting environmentally friendly components for solid rocket propellants at the early stages of design." Eastern-European Journal of Enterprise Technologies 6, no. 6 (114) (2021): 6–14. http://dx.doi.org/10.15587/1729-4061.2021.247233.

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This paper has investigated the possibility to theoretically calculate a value of the specific impulse for highly energetic compositions using only two parameters – the heat of the reaction and the number of moles of gaseous decomposition reaction products. Specific impulse is one of the most important energetic characteristics of rocket propellant. It demonstrates the level of achieving the value of engine thrust and propellant utilization efficiency. Determining the specific impulse experimentally is a complex task that requires meeting special conditions. For the stage of synthesis of new p

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36

Petrov, Nikolay, and Tamara Antonova. "Increasing the specific impulse of the ion engine by zone engineering of the solid-state field cathode." Proceedings of the Russian higher school Academy of sciences, no. 4 (January 20, 2021): 41–50. http://dx.doi.org/10.17212/1727-2769-2020-4-41-50.

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With the rapid development of space technology, the scale of human space exploration is expanding significantly. However, the growing demand for deep space travel cannot be met with conventional chemical engines. Thus, the need for new mechanisms for providing jet thrust, including electric motors, becomes clear. Electric propulsion technology has significant advantages over traditional chemical engines in deep space flight due to its characteristics such as high specific impulse, small size, long service life. A negative feature of electric motors can be called low thrust, however, firstly, i

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Eisen, Nachum E., and Alon Gany. "Theoretical Performance Evaluation of a Marine Solid Propellant Water-Breathing Ramjet Propulsor." Journal of Marine Science and Engineering 8, no. 1 (2019): 8. http://dx.doi.org/10.3390/jmse8010008.

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This work analyzes and presents theoretical performance of a marine water-breathing ramjet propulsor. A conceptual scheme of the motor is shown, the equation of thrust is presented, and the dependence on cruise velocity and depth are discussed. Different propellant compositions, representing a wide variety of formulations suitable for propelling a water-breathing ramjet, are investigated. The theoretical results reveal that the specific impulse of a water-breathing ramjet can increase by as much as 30% compared to a standard rocket, when using a conventional hydroxyl terminated polybutadiene (

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Mahammadsalman Warimani, Muhammad Hanafi Azami, Sher Afghan Khan, et al. "Analytical Assessment of Blended Fuels for Pulse Detonation Engine Performance." Journal of Advanced Research in Fluid Mechanics and Thermal Sciences 93, no. 2 (2022): 1–16. http://dx.doi.org/10.37934/arfmts.93.2.116.

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A pulse detonation engine (PDE) is possible to be a next-generation high-performance propulsion system in aerospace-related applications. To generate power or thrust, PDE uses repeated detonations. The current study evaluates the PDE performance with alternative and blended fuels in the Zeldovich–von Neumann–Doring (ZND) model. Parameters such as temperature ratio, pressure ratio, detonation velocity, and specific impulse were determined analytically for the fuels. The computed detonation parameters and specific impulse were compared with those available in NASA’s open-source program, Chemical

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Shi, Lisong, E. Fan, Hua Shen, Chih-Yung Wen, Shuai Shang, and Hongbo Hu. "Numerical Study of the Effects of Injection Conditions on Rotating Detonation Engine Propulsive Performance." Aerospace 10, no. 10 (2023): 879. http://dx.doi.org/10.3390/aerospace10100879.

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A three-dimensional upwind conservation element and solution element method (CESE) in cylindrical coordinates is first developed to effectively solve the unsteady reactive Euler equations governing a hydrogen–air rotating detonation engine (RDE) with coaxial structures. The effects of the annular width on the structure of the detonation front and the relationship between the thrust and mass flow rate are then investigated. Additionally, RDEs with various injection conditions are systematically analyzed regarding flow patterns and propulsion performance. The results reveal a positive correlatio

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Yachmenev, P. S., V. V. Fedyanin, and I. S. Vavilov. "DEVELOPMENT OF A THRUST MEASUREMENT STAND BASED ON THE AERODYNAMIC METHOD FOR ELECTRIC THRUSTERS OF SMALL SPACECRAFT." DYNAMICS OF SYSTEMS, MECHANISMS AND MACHINES 11, no. 2 (2023): 51–57. http://dx.doi.org/10.25206/2310-9793-2023-11-2-51-57.

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A traction measurement stand based on the aerodynamic method has been developed. To test the stand, a prototype electric rocket engine was used as a cold gas engine. The working medium was nitrogen gas. At a flow rate of 1.2 mg/s, the thrust value was from 0.674 to 0.736 mN, at a flow rate of 1.7 mg/s, the thrust value was 0.934-1 mN and at a flow rate of 2.6 mg/s, the thrust value was from 1.59 to 1.634 mN. The experimental specific thrust impulse was at a flow rate of 1.2 mg/s from 574 to 613 m/s, at a flow rate of 1.7 mg/s from 549 to 588 m/s and at a flow rate of 2.6 mg/s from 612 to 623 m

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Rezaei, Hadi, and Mohammad Reza Soltani. "An analytical and experimental study of a hybrid rocket motor." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 228, no. 13 (2014): 2475–86. http://dx.doi.org/10.1177/0954410013519432.

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The hybrid rocket motor is a kind of chemical propulsion system that has been recently given serious attention by various industries and research centers. The relative simplicity, safety and low cost of this motor, in comparison with other chemical propulsion motors, are the most important reasons for such interest. Moreover, throttle-ability and thrust variability on demand are additional advantages of this type of motor. In this paper, the result of an internal ballistic simulation of hybrid rocket motor in a zero-dimensional form is presented. Further to validate the code, an experimental s

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Genta, Giancarlo, and Dario Riccobono. "Optimization of Interplanetary Trajectory for Direct Fusion Drive Spacecraft." Journal of the British Interplanetary Society 76, no. 5 (2023): 170–77. http://dx.doi.org/10.59332/jbis-076-05-170.

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The Direct Fusion Drive (DFD) technology, which is being developed at present, will allow fast and affordable interplanetary travel. This is a result of the very high specific impulse and the low specific mass of DFD thrusters which outperform more conventional Nuclear Electric Propulsion (NEP), with which it shares the ability of providing a low (albeit higher than in the case of NEP) continuous thrust. It is well known that, to optimize the payload fraction, the thruster should operate in Variable Exhaust Velocity (VEV) mode and that the lower is the specific mass, the higher should be the m

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Genta, Giancarlo, and Dario Riccobono. "Optimization of Interplanetary Trajectory for Direct Fusion Drive Spacecraft." Journal of the British Interplanetary Society 76, no. 5 (2023): 170–77. http://dx.doi.org/10.59332/jbis-076-05-0170.

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The Direct Fusion Drive (DFD) technology, which is being developed at present, will allow fast and affordable interplanetary travel. This is a result of the very high specific impulse and the low specific mass of DFD thrusters which outperform more conventional Nuclear Electric Propulsion (NEP), with which it shares the ability of providing a low (albeit higher than in the case of NEP) continuous thrust. It is well known that, to optimize the payload fraction, the thruster should operate in Variable Exhaust Velocity (VEV) mode and that the lower is the specific mass, the higher should be the m

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Yao, Zhaohui, Shan Zhou, Tianlin Yang, and Yani Han. "Thermal Performance Characteristics of an 80-Ton Variable-Thrust Liquid Engine for Reusable Launch Rockets." Sustainability 15, no. 8 (2023): 6552. http://dx.doi.org/10.3390/su15086552.

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In this paper, an 80-ton thrust liquid rocket engine (hereinafter referred to as an LRE) with a gas generator cycle, a 5:1 thrust throttling ratio, and an integrated flow regulator/gas generator (hereinafter referred to as an IFRGG) is analyzed. This LRE can be used during the first stage of launching, second-stage and upper-stage space missions, and moon/mars low-orbit hovering and soft landing, and it can also be used with various near-space multipurpose flight vehicles. The thermal performance model of the variable-thrust LRE is established, the influence of the main LRE design parameters o

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Munro-O’Brien, Thomas F., Mohamed Ahmed, Andrea Lucca Fabris, and Charles N. Ryan. "Inter-Laboratory Characterisation of a Low-Power Channel-Less Hall-Effect Thruster: Performance Comparisons and Lessons Learnt." Aerospace 12, no. 7 (2025): 601. https://doi.org/10.3390/aerospace12070601.

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A collaborative inter-laboratory study was conducted to characterise the performance of the novel 250 W External Discharge Plasma Thruster (XPT) with a channel-less Hall effect-type thruster designed to address lifetime limitations and lower-power efficiency challenges in conventional Hall effect thrusters. This study aimed to validate performance measurements across different facilities and thrust stands, investigating potential facility effects on thrust characterisation. Performance testing was conducted both at the University of Surrey using a torsional thrust balance and at the University

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Qi, Rui, and ShiJie Xu. "Optimal Low-Thrust Transfers to Lunar L1 Halo Orbit Using Variable Specific Impulse Engine." Journal of Aerospace Engineering 28, no. 4 (2015): 04014096. http://dx.doi.org/10.1061/(asce)as.1943-5525.0000432.

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Semenov, V. L., V. Yu Aleksandrov, A. N. Prokhorov, K. Yu Arefyev, and S. V. Kruchkov. "Methodological Aspects of Determining Thrust of Irrotational Air-Breathing Jet Engines in Bench and Flight Tests." Proceedings of Higher Educational Institutions. Маchine Building, no. 11 (716) (November 2019): 86–97. http://dx.doi.org/10.18698/0536-1044-2019-11-86-97.

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This article examines methodological aspects of the indirect calculation of thrust characteristics of irrotational air-breathing jet engines using telemetry data that can be obtained during high-speed aircraft flight tests. Specific features of determining thrust characteristics during bench and flight tests are described. Mathematical models are developed for data analysis and calculation of the thrust and the specific impulse of an irrotational air-breathing jet engine by internal parameters, as well as its effective thrust in integration with a high-speed aircraft. The proposed approaches a

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Glascock, Matthew S., Joshua L. Rovey, and Kurt A. Polzin. "Impulse and Performance Measurements of Electric Solid Propellant in a Laboratory Electrothermal Ablation-Fed Pulsed Plasma Thruster." Aerospace 7, no. 6 (2020): 70. http://dx.doi.org/10.3390/aerospace7060070.

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Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. This behavior may enable the propellant to be used in multimode propulsion systems utilizing the ablative pulsed plasma thruster. The performance of an electric solid propellant operating in an electrothermal ablation-fed pulsed plasma thruster was investigated using an inverted pendulum micro-newton thrust stand. The impulse bit and specific impulse of the device using the electric solid propellant we

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Arzhannikov, Andrey, and Alexey Beklemishev. "An Electro-Jet Rocket Engine With Big Thrust At Helical Corrugated Magnetic Field." Siberian Journal of Physics 11, no. 1 (2016): 107–18. http://dx.doi.org/10.54362/1818-7919-2016-11-1-107-118.

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A fundamentally new electro-jet rocket engine having a big thrust with a high specific impulse is described in this paper. The acceleration mechanism of magnetized plasma along the axis of a cylindrical chamber with a helical corrugated magnetic field is put in the basis of such engine. The plasma acceleration is achieved during its drift motion by applying a radial electric field. The analytical description of the plasma motion process gives a visual representation of how the diamagnetic forces provide the process of the continuous acceleration of plasma ions along the axis of the helical cor

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Liu, Bendong, Xinrui Li, Jiahui Yang, and Guohua Gao. "Recent Advances in MEMS-Based Microthrusters." Micromachines 10, no. 12 (2019): 818. http://dx.doi.org/10.3390/mi10120818.

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With the development of micro/nano satellites and formation flying, more advanced spatial propulsion technology is required. In this paper, a review of microthrusters developments that based on micro electromechanical systems (MEMS) technology adopted in microthrusters is summarized. The microthrusters in previous research are classified and summarized according to the types of propellants and the working principles they utilized. The structure and the performance including the thrust, the impulse and the specific impulse of various microthrusters are compared. In addition, the advantages and

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